Ceramic materials are often used in high temperature applications such as the hot combustion gas path components of a gas turbine engine. Monolithic ceramic materials generally exhibit higher operating temperature limits than do metals, however they lack the toughness and tensile load carrying capabilities required for most structural applications. Ceramic matrix composite (CMC) materials are known to provide a combination of high temperature capability, strength and toughness.
FIG. 1 is a cross-sectional view of a prior art component, specifically a stationary airfoil or vane 10 for a gas turbine engine that is formed using ceramic materials. Vane 10 includes a layer of a very high temperature ceramic insulating material 12 disposed over a CMC structural member 14, such as is described in U.S. Pat. No. 6,197,424, incorporated by reference herein in its entirety. The CMC structural member 14 defines a plurality of passages 16 for directing a flow of cooling air. Internal ribs or spars 18 are formed to stiffen the structure. One or both radial ends of the airfoil 10 may be supported in a platform (not illustrated) of a gas turbine engine. A layer of adhesive 20 may be used to join the insulating material 12 to the CMC structural member 14. The CMC structural member 14 may typically be formed by laying up a plurality of plies of material in stacked planes that are parallel to the exterior surface of the member 14. A predetermined number of such plies of material are used to achieve a desired thickness dimension (perpendicular to the exterior surface) in the CMC structural member 14. The plies of material are thus wrapped around the leading edge portion 22 of the airfoil 10. Interlaminar stresses between adjacent plies can result from internal pressure in the cooling air passages 16, from thermal gradients across the CMC material, and from operating loads imposed on the airfoil 10.